Electric power system architecture and fault tolerant VTOL aircraft using same

ABSTRACT

A power system with a reliability enhancing battery architecture for electric motors adapted for use in an aerial vehicle. Individual batteries may be used to power a subset two or more motors in systems with six or more motors, for example. Each motor may be powered may be powered by two or more subsets of batteries, allowing accommodation for motor failure. With a failed motor in a vertical take-off or landing mode, power may be diverted to other motors to continue proper attitude control, and to provide sufficient thrust. With a failed motor a second motor offset from the failed motor may be powered down to facilitate attitude control.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Patent ApplicationNo. 62/678,275 to Bevirt et al., filed May 31, 2018, which is herebyincorporated by reference in its entirety.

FIELD OF THE INVENTION

This invention relates to electric powered flight, namely a power systemfor electric motors used on aerial vehicles.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1A-1D are of a VTOL aircraft in a hover configuration according tosome embodiments of the present invention.

FIGS. 1E-1H are of a VTOL aircraft in a forward flight configurationaccording to some embodiments of the present invention.

FIGS. 1I-1K are of a VTOL aircraft transitioning from a forward flightconfiguration to a vertical take-off and landing configuration accordingto some embodiments of the present invention.

FIG. 2A is a layout of a flight system with a ring architectureaccording to some embodiments of the present invention.

FIG. 2B is a layout identifying motor locations for the ringarchitecture according to some embodiments of the present invention.

FIG. 2C is a layout of battery locations according to some embodimentsof the present invention.

FIG. 3 is a motor power chart according to some embodiments of thepresent invention.

FIG. 4 is a failure scenario layout according to some embodiments of thepresent invention.

FIG. 5 is a failure compensation layout according to some embodiments ofthe present invention.

FIG. 6 is a failure compensation layout according to some embodiments ofthe present invention.

FIG. 7 is a power architecture layout according to some embodiments ofthe present invention.

FIG. 8 is a battery discharge chart according to some embodiments of thepresent invention.

FIG. 9 is a flight control system architecture layout according to someembodiments of the present invention.

FIG. 10 illustrates flight control software architecture according tosome embodiments of the present invention.

FIG. 11A is a layout of a flight power system with a doubletarchitecture according to some embodiments of the present invention.

FIG. 11B is a layout of a flight power system with a doubletarchitecture according to some embodiments of the present invention.

FIG. 11C is a layout of a flight power system with a doubletarchitecture with a motor failure according to some embodiments of thepresent invention.

FIG. 11D is a layout of a flight power system with a doubletarchitecture with a battery failure according to some embodiments of thepresent invention.

FIG. 12 is a layout of a flight power system with a hexagramarchitecture according to some embodiments of the present invention.

FIG. 13 is a layout of a flight power system with a star architectureaccording to some embodiments of the present invention.

FIG. 14 is a layout of a flight power system with a mesh architectureaccording to some embodiments of the present invention.

FIGS. 15A-C presents information regarding battery failure operationsaccording to some embodiments of the present invention.

SUMMARY

A power system with a reliability enhancing power system architecturefor electric motors adapted for use in an aerial vehicle. Individualbatteries may be used to power a subset of two or more motors in systemswith six or more motors, for example. Each motor may be powered by twoor more subsets of batteries, allowing accommodation for motor failure.Each motor may have two or more sets of windings, with each windingpowered by a different battery. With a failed winding, failed battery,or failed motor in a forward flight or a vertical take-off and landingmode, power routing may be automatically altered to continue properattitude control, and to provide sufficient thrust. With a failed motora second motor offset from the failed motor may be powered down tofacilitate attitude control.

DETAILED DESCRIPTION

In some aspects, an aerial vehicle may use bladed propellers powered byelectric motors to provide thrust during take-off. The propeller/motorunits may be referred to as propulsion assemblies. In some aspects, thewings of the aerial vehicle may rotate, with the leading edges facingupwards, such that the propellers provide vertical thrust for take-offand landing. In some aspects, the motor driven propeller units on thewings may themselves rotate relative to a fixed wing, such that thepropellers provide vertical thrust for take-off and landing. Therotation of the motor driven propeller units may allow for directionalchange of thrust by rotating both the propeller and the electric motor,thus not requiring any gimbaling, or other method, of torque drivearound or through a rotating joint.

In some aspects, aerial vehicles according to embodiments of the presentinvention take off from the ground with vertical thrust from rotorassemblies that have deployed into a vertical configuration. As theaerial vehicle begins to gain altitude, the rotor assemblies may beginto be tilted forward in order to begin forward acceleration. As theaerial vehicle gains forward speed, airflow over the wings results inlift, such that the rotors become less important and then unnecessaryfor maintaining altitude using vertical thrust. Once the aerial vehiclehas reached sufficient forward speed, some or all of the blades used forproviding vertical thrust during take-off may be stowed along theirnacelles. In some aspects, all propulsion assemblies used for verticaltake-off and landing are also used during forward flight. The nacellesupporting the propulsion assemblies may have recesses such that theblades may nest into the recesses, greatly reducing the drag of thedisengaged rotor assemblies.

After take-off, the aerial vehicle will begin a transition to forwardflight by articulating the propellers from a vertical thrust orientationto a position which includes a horizontal thrust element. As the aerialvehicle begins to move forward with speed, lift will be generated by thewings, thus requiring less vertical thrust form the rotors. As thepropellers are articulated further towards the forward flight,horizontal thrust, configuration, the aerial vehicle gains more speed.

In a first vertical configuration according to some embodiments of thepresent invention, as seen in a vertical take-off configuration in FIGS.1A through 1D, an aerial vehicle 200 uses fixed wings 202, 203, whichmay be forward swept wings, with propulsion assemblies of the same ordifferent types adapted for both vertical take-off and landing and forforward flight. In this configuration, the propulsion assemblies arepositioned for vertical thrusting. The aircraft body 201 supports a leftwing 202 and a right wing 203. Motor driven propulsion assemblies 206along the wings may include electric motors and propellers which areadapted to articulate from a forward flight configuration to a verticalconfiguration using deployment mechanisms which may reside in thenacelle body, and which deploy the motor and propeller while all or mostof the nacelle remains in place attached to the wing. In some aspects,the propeller blades may stow and nest into the nacelle body. The motordriven rotor assemblies 207 at the wing tips may deploy from a forwardflight configuration to a vertical take-off and landing configurationalong a pivot axis wherein the nacelle and the electric motor andpropeller deploy in unison. Although illustrated with one mid-spanpropulsion assembly and one wingtip propulsion assembly per wing, insome aspects more mid-span propulsion assemblies may be present.

The aircraft body 201 extends rearward is also attached to raised rearstabilizers 204. The rear stabilizers have rear propulsion assemblies205 attached thereto. The motor driven rotor assemblies 205 at the tipsof the rear stabilizers may deploy from a forward flight configurationto a vertical take-off and landing configuration along a pivot axiswherein the nacelle and the electric motor and propeller deploy inunison.

As seen in top view in FIG. 1D, the propulsion assemblies are positionedat different distances from the aircraft center of mass, in two axes.Attitude control during vertical take-off and landing may be manipulatedby varying the thrust at each of the propulsion assembly locations. Inthe circumstance of a motor failure during vertical take-off or landing,and especially a motor failure at the wing outboard propulsion assembly,the attitude of the aircraft may be maintained by implemented faulttolerance strategies described herein.

The aerial vehicle 200 is seen with two passenger seats side by side, aswell as landing gear under the body 201. Although two passenger seatsare illustrated, other numbers of passengers may be accommodated indiffering embodiments of the present invention.

FIGS. 1E through 1H illustrates the aerial vehicle 200 in a forwardflight configuration. In this configuration, the propulsion assembliesare positioned to provide forward thrust during horizontal flight. Asseen in FIG. 1H, the centers of mass of the motors and of the propellersmay be forward of the leading edge of the wings in the forward flightconfiguration. As seen in FIG. 1G, the propulsion assemblies 205 on therear stabilizers 204 may be at a different elevation that the propulsionassemblies 206, 207 on the wings. In the circumstance of a motor failureduring forward flight, the attitude of the aircraft may be maintained byimplementing fault tolerance strategies described herein.

In some aspects, all or a portion of the wing mounted propulsionassemblies may be adapted to be used in a forward flight configuration,while other wing mounted propellers may be adapted to be fully stowedduring regular, forward, flight. The aerial vehicle 200 may have twopropulsion assemblies on the right wing 203 and two propulsionassemblies on the left wing 202. The inboard propulsion assemblies oneach wing may have wing mounted rotors that are adapted to flip up intoa deployed position for vertical take-off and landing, to be moved backtowards a stowed position during transition to forward flight, and thento have their blades stowed, and nested, during forward flight. Theoutboard propulsion assembly 207 may pivot in unison from a horizontalto a vertical thrust configuration.

Similarly, the each rear stabilizer 204 may have propulsion assembliesmounted to it, both of which are adapted to be used during verticaltake-off and landing, and transition, modes. In some aspects, all of thepropulsion assemblies designs are the same, with a subset used withtheir main blades for forward flight. In some aspects, all of thepropulsion assemblies designs are the same, with all propellers used forforward flight. In some aspects, there may be a different number ofpropulsion assemblies units mounted to the rear stabilizer 204.

The motors driving the wing mounted propulsion assemblies 206, 207 andthe motors driving the rear stabilizer mounted propulsion assemblies mayeach have two sets of windings. In some aspects, both winding sets arepowered during flight. In some aspects, each winding of the motor ispowered by a different battery circuit. In some aspects, each motor mayhave more than two sets of windings.

In some embodiments, the electric motors of the aerial vehicle arepowered by rechargeable batteries. The use of multiple batteries drivingone or more power busses enhances reliability, in the case of a singlebattery failure. In some embodiments, the batteries reside within thevehicle body on a rack with adjustable position such that the vehiclebalance may be adjusted depending upon the weight of the pilot. FIG. 2Aillustrates a battery location layout for a six battery system accordingto some embodiments of the present invention.

In some embodiments, as seen in FIG. 2A, a high reliability power system10 for an electrically powered vertical take-off and landing aircrafthas six motors and six batteries in a ring architecture. In thisexemplary configuration, there are six motors and six batteries. Each ofthe batteries provides power to two motors, and each motor receivespower from two batteries. FIG. 2B illustrates a layout of six motors ona VTOL aircraft in an exemplary embodiment using six propulsionassemblies and six batteries. FIG. 2C illustrates a layout of sixbatteries in a VTOL aircraft in an exemplary embodiment using sixpropulsion assemblies and six batteries. In an exemplary ringembodiment, there are six batteries and six motors. Each of the motorsis powered by two separate batteries. The disparate locations 30 of thebatteries also enhance the reliability and fault tolerance of the powersystem architecture. Each battery is powering two separate motors. Insome aspects, each of the motors is wound with two sets of windings, andeach set of windings receives power from a different battery. Asdiscussed below with regard to FIG. 7 , each of the six batteriessupplies two power inverters 31, for a total of 12 power inverters. Thenominal voltage of the batteries is 600V. Each of the six propulsionmotors has two sets of windings, with each motor powered by twoinverters, one for each set of windings. The two inverters powering asingle motor each are supplied power by different batteries.

In an exemplary six motor six battery embodiment 10, the first motor 11is coupled the sixth battery 26 and the first battery 21. The secondmotor 12 is coupled to the first battery 21 and the second battery 22.The third motor 13 is coupled to the second battery 22 and the thirdbattery 23. The fourth motor 14 is coupled to the third battery 23 andthe fourth battery 24. The fifth motor 15 is coupled to the fourthbattery 24 and the fifth battery 25. The sixth motor 16 is coupled tothe fifth battery 25 and the sixth battery 26. In a nominal operatingscenario, each battery splits its power distribution evenly between thetwo motors to which it is coupled, and each motor receives an equalamount of power from each battery to which it is coupled.

The fault tolerant aspect of the power system architecture according toembodiments of the present invention is adapted to withstand, andrespond to, at least the following failures: the failure of a battery;the failure of a motor; or the failure of a motor invertor.

FIG. 3 is a bar graph (with bar pairs for each operating mode) of thepower required for a single motor 40 in the six motor embodiment. Theblue vertical bars (on the left side of the bar pair for each mode)illustrate nominal (normal) operating power, per motor, for the fivedifferent flight phases: hover 41, vertical ascent 42, vertical descent43, cruise climb 44, and cruise 45. The hover, vertical ascent, andvertical descent modes are VTOL modes wherein the motors are rotated toa vertical thrust position as seen in FIGS. 1A-1D. The cruise climb andcruise phases are with the motors in a forward flight position, as seenin FIG. 1E-1H. The red vertical bars (on the right side of the bar pairfor each mode) represent emergency phase operation, as discussed below.

As seen in FIG. 3 , the illustrative embodiment of a six motor sixbattery ring architecture system runs about 60 kW per motor in a VTOLmode during nominal conditions. This 60 kW compares to approximately the150 kW maximum available power. In the case of a motor failure, however,more power may be diverted to remaining motors to maintain attitude andaltitude control, as discussed further below.

FIG. 4 illustrates a potential failure mode 60 wherein the first motorfails. As seen in the representation motor layout, the loss of the firstmotor 11 represents loss of thrust at the far port motor, which willhave a significant impact on the attitude of the aircraft. The flightcomputer may immediately sense at least two things: first, that themotor has quit drawing current; second, that there is a disruption tothe attitude of the aircraft. In order to maintain balance in theaircraft, the flight control computer will reduce power to the opposingmotor(s) as needed. In this example, as seen in FIG. 5 , the power tothe fourth motor 14 will be reduced. The loss of lift due to theshutdown of two motors requires that the remaining four motors take morepower and deliver more lift. FIG. 6 illustrates how the increased loaddemands in the second, third, fifth, and sixth motors are met bydistributing more power from the batteries. Looking again at FIG. 3 ,the red vertical bars illustrate the power delivery required with themotor failure and then the motor shutdown of the opposing motor. Thepower down of the fourth motor and the increase in power to the second,third, fifth, and sixth motors may take place simultaneously in someaspects. In some aspects, the power down of the fourth motor and theincrease in power to the second, third, fifth, and sixth motors may takeplace sequentially.

As seen in FIG. 6 , with the first motor 11 failed and the fourth motor14 powered down to balance the aircraft, the first battery 21 now onlydelivers power to the second motor 12. Similarly, the third battery onlydelivers power to the third motor, the fourth battery only deliverspower to the fifth motor, and the sixth battery only delivers power tothe sixth motor. The second battery delivers power to both the secondand third motors, and the fifth battery delivers power to the fifth andsixth motors. Although illustrated as having the fourth motor runningdown to 0% power, in some aspects the cross motor may be run at a lowlevel, in the range of 0-20% of nominal power, for example. As the firstand sixth battery are only providing to a single motor, and as the thirdand fourth battery are primarily only delivering power to a singlemotor, these batteries will provide more current 61 to their respectivewindings in the second, third, fifth, and sixth motors. The second andfifth batteries will split evenly between their adjacent motors. In thefailure scenario illustrated in FIG. 6 , each battery may be putting outthe same amount of power, but two batteries are splitting their powerdelivery, and four batteries are providing (or substantially providing)power to only a single motor. The increased load demand of the motors inthis emergency mode is shared through the battery architecture toutilize the available energy onboard the aircraft. Although one motorhas been disabled and a second motor has been powered down toaccommodate attitude control concerns, each battery is still being usedand delivering power.

In some embodiments, the vertical take-off and landing aircraft has anautonomous attitude control system adapted to withstand a power linkfailure, or complete motor failure, in a multi-battery system by loadsharing to better equate battery discharge levels. In some aspects, eachmotor is driven on multiple complementary winding sets, with eachwinding set using a different load link and being driven by a differentbattery. FIG. 7 is an illustrative embodiment of the electrical systempower architecture for a six motor six battery aircraft. Each of the sixbatteries 251 supplies two power inverters, for a total of 12 powerinverters 252. The nominal voltage of the batteries is 600V. Each of thesix propulsion motors 253 has two sets of windings, with each motorpowered by two inverters, one for each set of windings. The twoinverters powering a single motor each are supplied power by differentbatteries. In addition to supplying power to the motor inverters, thebattery also supplies power to the rotor deployment mechanisms 254(nacelle tilt actuators) which are used to deploy and stow the rotorsduring various flight modes (vertical take-off and landingconfiguration, forward flight configuration, and transition between).

A flight computer 255 monitors the current from each of the twelve motorinverters 252 which are supplying power to the twelve winding sets inthe six motors 253. The flight computer 255 may also control the motorcurrent supplied to each of the 12 sets of windings of the six motors.In some embodiments, the batteries 251 also supply power to the bladepitch motors and position encoders of the variable pitch propellers 256.The batteries also supply power to control surface actuators 257 used toposition various control surfaces on the airplane. The blade pitchmotors and the control surface actuators 257 may receive power runthrough a DC-DC converter 258, stepping the voltage down from 600V to160V, for example. A suite of avionics 259 may also be coupled to theflight computer. A battery charger 250 may be used to recharge thebatteries 251, and the battery charger may be external to the aircraftand ground based.

In the case of a failure, such as the failure of a motor, or of a powerlink to a motor, the compensations to power distribution to the variousmotors from the various batteries, as described above, may be doneautonomously and onboard the aircraft. The compensations may be donewithout needing input from the pilot, for example.

In another failure scenario, a single winding on a motor may fail. Insuch a scenario, the opposing motor may be powered down somewhat whilethe motor with a sole remaining winding may be powered up somewhat. Thepower supplied by the batteries may be moderated to even out thedischarge of the various batteries. In yet another failure scenario, abattery may fail. In that case the cross motor may be reduced 10-20%,with the sole battery remaining on the motor with the failedbattery/inverter providing extra power, and differential power along thering used to spread the battery discharge. In the case of a completelyfailed battery in the ring architecture, which would result in twomotors each having one winding set go unpowered, the remaining windingset in each of the adjacent motors would take increased power from thatwinding set's battery, and there would be differentially adjusted poweraround the ring in order to best equalize the battery discharge rates.The cross motor would be partially powered down to maintain appropriatedischarge rates.

FIG. 8 illustrates four flight modes and a bar chart 235 of the batterydischarge rates for each flight mode. The vertical axis in the bar chartis the battery discharge rate C. The battery discharge rate is anormalized coefficient wherein a 1 C discharge rate would discharge thebattery in one hour. A 2 C discharge the battery in 30 minutes, a 3 Cdischarge rate would discharge the battery in 20 minutes, and so on. Amaximum peak discharge rate 236, which is about 5 C in this exemplaryembodiment, may be set by the limitations of the battery chemistry. Thenominal flight modes are hover mode 232, transition mode 233, and cruisemode 234. The cruise discharge rate 240 may be approximately 1 C. As theaircraft approaches for landing, the aircraft will change to atransition mode 233, which may have a transition discharge rate 239 ofapproximately 2 C. The aircraft will then go into hover mode 232 as itlands, which may have a discharge rate 238 of approximately 2.5 C. Inthe case of a failed motor, the aircraft may go into an emergency hovermode 231, wherein a cross motor may be powered down to achieve attitudestability. The emergency hover mode discharge rate 237 may be over 3 C.

In an exemplary embodiment, the maximum gross take-off weight (MGTOW)may be 4200 pounds. The discharge rates are out of ground effect (OGE),with a total energy storage of all batteries of 150 kWh. In the case ofan emergency landing in the emergency hover mode 231, the anticipatedtime using the high discharge rate at the emergency hover mode dischargerate 237 is approximately 1 minute.

FIG. 9 illustrates a flight control system architecture for a highreliability electric powered aircraft according to some embodiments ofthe present invention. In an exemplary embodiment, the flight computer111 of the control system receives flight commands 114 from the missioncomputer 112 and from the pilot 113. The flight computer may alsoreceive inputs from a flight critical sensor suite 110. The flightcritical sensors may be triply redundant. The flight computer may betriply redundant. The system may include a voting bridge 116 on eachactuator 115. FIG. 10 illustrates the flight control softwarearchitecture according to some embodiments of the present invention.

In some embodiments of the present invention, other battery and motorarchitectures may be used which further enhance the fault tolerance ofthe system. In some aspects, as seen in FIG. 11A, a doublet architecture120 is used which uses four batteries to the electric motors of sixpropulsion assemblies; a left wing tip propulsion assembly motor 121, aleft wing propulsion assembly motor 122, a right wing propulsionassembly motor 123, a right wing tip propulsion assembly motor 124, aleft rear propulsion assembly motor 125, and a right rear propulsionassembly motor 126. In the doublet architecture, each battery providespower to one or motors on each side of the aircraft longitudinalcenterline. By linking a battery that powers the furthest outboard to amotor on the other side of the center line of the aircraft, a batteryfailure then has its effect more spread out across the aircraft,reducing the amount of attitude offset due to the battery failure. Inthe case of a motor failure at the first motor 121, for example, theremay still be an instantaneous reduction in power to the fourth motor tocompensate for the failure. But the compensation regime for powersharing in a doublet architecture using the remaining motors will allowfor lower inverter loads in an inverter optimized system as compared tothe ring architecture that was disclosed above. Also, the compensationregime for power sharing in a doublet architecture using the remainingmotors will allow for lower battery loads in a battery optimized systemas compared to the ring architecture.

FIG. 11B illustrates a nominal operating condition for the doubletarchitecture 120 wherein each of the four batteries 131, 132, 133 and134 provides 35 KW to one winding of three different motors, for a totalof 105 kW delivered per battery, and a total of 70 kW received permotor, for a total delivery of 420 kW. Each motor is receiving powerfrom three batteries.

FIG. 11C illustrates a motor failure condition, in this exemplary casethe motor 121 of the left wing tip propulsion assembly. As illustrated,the motor 124 on the right wing tip has been unpowered and is no longerdrawing any power, in order to offset the loss of the left wing tipmotor. Each of the batteries is now powering two motors instead of theprior three, and each motor is receiving power from two batteries,instead of the prior three. Each of the batteries is able to run at thesame power output level, and each of the motor windings, and theirassociated inverters, are also able to run at the same power level.

FIG. 11D illustrates a battery failure condition, in this exemplary casethe first battery 131. In this circumstance, each remaining batteryprovides the same power output level, although the different motors runat different power levels in order to balance the thrust generated oneach side of the aircraft longitudinal centerline.

FIG. 12 illustrates a six battery six motor hexagram architectureaccording to some embodiments of the present invention. In the hexagramarchitecture illustrated in FIG. 12 , each of the six batteries powerstwo motors, as with the ring architecture. And each motor is powered bytwo batteries. However, the first battery provides power to the firstand third motor, the second battery provides power to the second andfourth motor, and so on. The hexagram architecture creates two separaterings encompassing the first, third, and sixth motors, and the second,fourth, and fifth motors. By linking a battery that powers the furthestoutboard to a motor on the other side of the center line of theaircraft, a battery failure then has its effect more spread out acrossthe aircraft, reducing the amount of attitude offset due to the batteryfailure. In the case of a motor failure at the first motor, for example,there may still be an instantaneous reduction in power to the fourthmotor to compensate for the failure. But the compensation regime forpower sharing in a hexagonal architecture using the remaining motorswill allow for lower inverter loads in an inverter optimized system ascompared to the ring architecture. Also, the compensation regime forpower sharing in a hexagonal architecture using the remaining motorswill allow for lower battery loads in a battery optimized system ascompared to the ring architecture. FIGS. 15A to 15C illustrate themaximum loads in the inverters, batteries, and motors forinverter-optimized, battery-optimized, and motor-optimized solutions forthe various motor-battery architectures described herein during abattery failure. The hexagram architecture is indicated with a symbol,as opposed to a name like the other architectures, in FIGS. 15A to 15C.

FIGS. 13 and 14 illustrate six motor four battery systems according tosome embodiments of the present invention. FIG. 13 illustrates a stararchitecture using four batteries to power six motors. Each battery iscoupled to three motors. FIG. 14 illustrates a mesh architecture withfour batteries and six motors.

FIGS. 15A, 15B, and 15C illustrate the maximum loads in the inverters,batteries, and motors, respectively, for inverter-optimized,battery-optimized, and motor-optimized solutions for the variousmotor-battery architectures described herein during a motor failure. Thehexagram architecture is indicated with a symbol, as opposed to a namelike the other architectures. As illustrated, the hexagram architecturegives the best solution when reviewed with regard to all optimizations(inverter-optimized, battery-optimized, and motor-optimized).

As evident from the above description, a wide variety of embodiments maybe configured from the description given herein and additionaladvantages and modifications will readily occur to those skilled in theart. The invention in its broader aspects is, therefore, not limited tothe specific details and illustrative examples shown and described. Theembodiments described herein may include physical structures, as well asmethods of use. Accordingly, departures from such details may be madewithout departing from the spirit or scope of the applicant's generalinvention.

What is claimed is:
 1. An electrically powered vertical take-off andlanding aircraft, said aircraft comprising: a left inboard propulsionassembly and a left outboard propulsion assembly. each of the leftinboard and left outboard propulsion assemblies comprising an electricmotor; a right inboard propulsion assembly and a right outboardpropulsion assembly,. each of the right inboard and right outboardpropulsion assemblies comprising an electric motor; a plurality ofbatteries, each of said plurality of batteries coupled to two or more ofsaid electric motors; and a flight control system adapted to: operatethe propulsion assemblies in a vertical thrust generating mode; detectfailure of a particular propulsion assembly; reduce power to acorresponding propulsion assembly, the corresponding propulsion assemblybeing symmetrically located on an opposite side of a longitudinalcenterline of said aircraft relative to the particular propulsionassembly; and increase the power delivered to some or all of theremaining propulsion assemblies in order to maintain necessary thrust.2. The aircraft of claim 1 further comprising a plurality of inverters,wherein each of said batteries is coupled to each of said electricmotors through an inverter.
 3. The aircraft of claim 1 wherein the leftinboard propulsion assembly is offset from the left outboard propulsionassembly along the direction of the longitudinal centerline of saidaircraft and the right inboard propulsion assembly is offset from theright outboard propulsion assembly along the direction of thelongitudinal centerline of said aircraft.
 4. The aircraft of claim 2wherein the flight control system is adapted to autonomously adjust thepower delivered by said batteries to said electric motors in the eventof a motor failure to maintain a desired aircraft attitude.
 5. Theaircraft of claim 1 wherein the flight control system is adapted toautonomously adjust the power delivered by said batteries to saidelectric motors in the event of a battery failure to maintain a desiredaircraft attitude.
 6. The aircraft of claim 2 wherein the flight controlsystem is adapted to autonomously adjust the power delivered by saidbatteries to said electric motors in the event of a battery failure tomaintain a desired aircraft attitude.
 7. The aircraft of claim 1 whereineach of said electric motors comprises a plurality of motor windingcircuits, and wherein each of said winding circuits in said plurality ofmotor winding circuits in a motor are coupled to a different battery. 8.The aircraft of claim 2 wherein each of said electric motors comprises aplurality of motor winding circuits, and wherein each of said windingcircuits in said plurality of motor winding circuits in a motor arecoupled to a different battery.
 9. The aircraft of claim 4 wherein eachof said electric motors comprises a plurality of motor winding circuits,and wherein each of said winding circuits in said plurality of motorwinding circuits in a motor are coupled to a different battery.
 10. Theaircraft of claim 6 wherein each of said electric motors comprises aplurality of motor winding circuits, and wherein each of said windingcircuits in said plurality of motor winding circuits in a motor arecoupled to a different battery.
 11. The aircraft of claim 1 wherein theparticular propulsion assembly is coupled to a first battery on a sameside of the longitudinal centerline of the aircraft as the particularpropulsion assembly and is coupled to a second battery on the oppositeside of the longitudinal centerline of the aircraft from the particularpropulsion assembly. and the flight control system is adapted to divertpower from the first battery and the second battery from the particularpropulsion assembly to some or all of the remaining propulsionassemblies.
 12. The aircraft of claim 2 wherein the particularpropulsion assembly is coupled to a first battery on a same side of thelongitudinal centerline of the aircraft as the particular propulsionassembly and is coupled to a second battery on the opposite side of thelongitudinal centerline of the aircraft from the particular propulsionassembly and the flight control system is adapted to divert power fromthe first battery and the second battery from the particular propulsionassembly to some or all of the remaining propulsion assemblies.
 13. Theaircraft of claim 7 wherein the particular propulsion assembly iscoupled to a first battery on a same side of the longitudinal centerlineof the aircraft as the particular propulsion assembly and is coupled toa second battery on the opposite side of the longitudinal centerline ofthe aircraft from the particular propulsion assembly, and the flightcontrol system is adapted to divert power from the first battery and thesecond battery from the particular propulsion assembly to some or all ofthe remaining propulsion assemblies.
 14. The aircraft of claim 1 whereinthe flight control system is adapted to detect failure of a particularpropulsion assembly by: detecting that the electric motor in theparticular propulsion assembly has stopped drawing current.
 15. Theaircraft of claim 1 wherein the flight control system is adapted toreduce power to the corresponding propulsion assembly by: reducing powerto be within a range of 0-20% of nominal power.
 16. An electricallypowered vertical take-off and landing aircraft, said aircraftcomprising: a left inboard propulsion assembly and a left outboardpropulsion assembly, the left inboard propulsion assembly being offsetfrom the left outboard propulsion assembly along the direction of alongitudinal centerline of said aircraft, each of the left inboard andleft outboard propulsion assemblies comprising an electric motor; aright inboard propulsion assembly and a right outboard propulsionassembly, the right inboard propulsion assembly being offset from theright outboard propulsion assembly along the direction of thelongitudinal centerline of said aircraft, each of the right inboard andright outboard. propulsion assemblies comprising an electric motor; aplurality of batteries, each of said plurality of batteries coupled totwo or more of said electric motors; and a flight control system adaptedto: operate the propulsion assemblies in a vertical thrust generatingmode; detect failure of a particular propulsion assembly of the left andright inboard and outboard propulsion assemblies; reduce power to acorresponding propulsion assembly, the corresponding propulsion assemblybeing symmetrically located on an opposite side of the longitudinalcenterline of said aircraft relative to the particular propulsionassembly; and increase the power delivered to some or all of theremaining propulsion assemblies in order to maintain necessary thrust.17. The aircraft of claim 16 further comprising a plurality ofinverters, wherein each of said batteries is coupled to each of saidelectric motors through an inverter.
 18. The aircraft of claim 16further wherein the flight control system is further adapted toautonomously adjust the power delivered by said batteries to saidelectric motors in the event of a battery failure to maintain a desiredaircraft attitude.
 19. The aircraft of claim 16 wherein the particularpropulsion assembly is coupled to a first battery on a same side of thelongitudinal centerline of the aircraft as the particular propulsionassembly and is coupled to a second battery on the opposite side of thelongitudinal centerline of the aircraft from the particular propulsionassembly, and the flight control system is adapted to divert power fromthe first battery and the second battery from the particular propulsionassembly to some or all of the remaining propulsion assemblies.
 20. Theaircraft of claim 16 wherein the flight control system is adapted toreduce power to the corresponding propulsion assembly by reducing powerto he within a range of 0-20% of nominal power.